Fan for gas turbine unit

ABSTRACT

A gas turbine unit, in particular for an aircraft turbojet engine, comprises a fan rotating coaxially in front of an axial flow compressor. The fan has an outer blading ring fixed to a rim attached to a hub of the fan through an inner blading ring, and the inner blades are pivotally attached, at the root and tip ends respectively, to the hub and to the rim about pivot axes parallel to the axis of rotation of the hub. The inner blades are all inclined in the same direction and at the same angle in relation to the hub circumference. The rim can thus expand freely under the effect of the centrifugal forces applied to it in operation by the outer blading, without local distortion and without the application of any bending stresses to the inner blades.

I United States Patent [1 1 Bouiller et al.

[ Oct. 30, 1973 FAN FOR GAS TURBINE UNIT [73] Assignee: SocieteNationale DEtude Et De Construction De Moteurs DAviation, Paris, France22 Filed: June 17, 1971 21 Appl. No.: 154,070

30 Foreign Application Priority Data June 22, 1970 France 7022919 [52]US. Cl 417/408, 415/77, 416/157, 1

[56] References Cited UNITED STATES PATENTS Krebs 60/226 R Stirling 41'577 Bauger et al 417/79 X Primary ExaminerC. J. Husar AttorneyWilliam J.Daniel 57 ABSTRACT A gas turbine unit, in particular for an aircraftturbojet engine, comprises a fan rotating coaxially in front of an axialflow compressor. The fan has an outer blading ring fixed to a rimattached to a hub of the fan through an inner blading ring, and theinner blades are pivotally attached, at the root and tip endsrespectively, to the hub and to the rim about pivot axes parallel to theaxis of rotation of the hub. The inner blades are all inclined in thesame direction and at the same angle in relation to the hubcircumference. The rim can thus expand freely under the effect of thecentrifugal forces applied to it in operation by the outer blading,without local distortion and without the application of any bendingstresses to the inner blades.

6 Claims, 5 Drawing Figures Patented Oct. 30, 1973 5 Sheets-Sheet 1Patented Oct. 30,1973 3,768,933

3 Sheets-Sheet 2 Patented Oct. 30,1973 Y 3,768,933

1 3 Sheets-Sheet 5 FAN FOR GAS TURBINE UNIT This invention relates togas turbine units comprising an axial flow compressor and a fan. Itrelates in particular to aircraft by-pass type gas turbine jet enginesbut is equally applicable to an industrial gaz turbine unit.

The chief purpose of a fan in such an aircraft engine is to supply anairflow, known as the secondary airflow, which passes around theexterior of the compressor casing in order to supply the engine withby-pass air. The central part of the fan is located upstream of thecompressor and therefore generally supplies thelatter with air which isalready slightly compressed.

It is well-known, in particular in order to facilitate or renderpossible the attachment of fan blading comprising a large number ofblades to a relatively small diameter hub, to split the blading intoouter and inner bladerings which are separated by a rim. The roots ofthe outer blades are fixed to the rim and the latter is attached to thehub of the fan through a smaller number of blades constituting the innerblading. The rim is, of course, located as an extension of thecompressor casing, in order that it does not disturb the flow.

In accordance with the present invention, the blades of the innerblading are articulated at root and tip respectively to the hub and tothe rim, about axes parallel to that of the hub, and are similarlyinclined in relation to the fan assembly, that is to say that the planescontaining the two axes of articulation of one and the same blade are,in the rest condition, tangential to an imaginary cylinder coaxial withthe hub.

. In operation, the blades of the inner blading transmit the hubrotational torque to the rim. The latter extends elastically under theeffect of the centrifugal forces applied to it by the outer blading,without applying any bending stress to the inner blades, which alignthemselves by pivoting about their articulation axes and withoutproducing any local distortion of the rim. Due to the fact that theinner blades serve exclusively to transmit the torque, they operatesubstantially in tension and are not submitted to any other mechanicalstress so that both the number and the mass of these blades can bereduced.

In accordance with one preferred feature of the invention, the innerblading is set to give a transparency condition, that is to say so thatin design conditions they produce no aerodynamic effect. As a result,the inner blades do not experience any aerodynamic stress at full loadand this enables the mechanical strength, and therefore the mass ofthese blades, to be still further reduced.

This preferred feature furthermore has the advantage that the innerblading has no influence upon the airflow entering the compressor, atany rate in the neighbourhood of the design point,.so that the controlof the engine can in particular be simplified as the speed of thecompressor can be regulated independently of that of the fan.

The description which now follows and which relates to the accompanyingdrawings is given solely by way of non-limitative example. It detailsthe advantages of the invention and how they are achieved, and featurescontained both in the present text and in the drawings fall within thescope of the invention. In the drawings FIG. 1 is a half-axial sectionthrough the forward part of a by-pass type gas turbine jet engine withan axial compressor and fan, which embodies improvements in accordancewith the invention FIG. 2 is a velocity diagram illustrating the settingof inner blading of the fan FIG. 3 is a partial perspective view, to alarger scale, illustrating the assembly of the fan blades; and

FIG. 4 isa schematic transverse sectional view, on a small scale,illustrating the operation of the fan;

FIG. 5 is a diagrammaticillustration of a jet engine showing the generalenvironment of the invention.

The jet engine illustrated diagrammatically in FIG. 5, the forward partof which is shown in F IG. 1, is of the three-spool type comprising afan 1 located upstream of a low' pressure axial flow compressor 2 whichis followed by a high-pressure axial flow compressor 2a. The fan 1, thecompressor 2 and the high-pressure compressor 2a are respectivelydriven, in a conventional manner, by three turbines la, lb, 10 arrangedand in series flow downstream of a combustion chamber 2b which issupplied with combustion air by the compressors 2, 2a, the gases leavingthe last turbine 1a being discharged through a nozzle 2c to form apropulsion jet. In the drawings, the hub 3 of the fan 1 is shown, thisbeing fixed to a shaft 4 which is rotated by the final turbine la, andalso shown is the rotor 5 of the low-pressure compressor 2 which isfixed to a shaft 4a coaxial to the shaft 4 and rotated by thecorresponding turbine lb. The reference 7 indicates the forward part ofthe compressor casing and 8 the forward part of the external fairing ornacelle of the engine.

During operation of the engine, the air entering at 9, at the front endof the engine, enters the fan 1, is then compressed by the compressor 2and recompressed by the high-pressure compressor 2a, whilst the airentering the peripheral portion of the fan 1, at 10, is compressed bythe latter and then flows into a diffuser 11a in an annular duct 11defined between a fairing 6 surrounding the casing 7 and the fairing 8,in order in the normal manner to supply the jet engine with by-pass air.

The turbine 1a which drives the fan 1 is designed to rotate at arelatively low speed so that the fan has a moderate peripheral velocityof the order of 300 m/sec., with a consequent very low noise level.However, in order to supply the duct 11 with air of sufficient pressure,the fan 1 must operate with a compression ratio of the order of 1.55. Inorder to achieve this compression ratio with blading of such lowperipheral velocity, it is necessary to use a large number of blades. Toavoid the difficulties of attaching such a large number of blades to thesmall diameter hub 3, and to avoid the obstruction which this number ofblades would create upstream of the compressor, the blading is splitinto two portions, namely outer blading 12 comprising a large number ofblades (45 in the present embodiment), the roots of which are fixed in arim 13 located upstream of the casing 7, and inner blading 14 with asmall number of blades (9 in the embodiment illustrated) articulated atthe blade roots to the hub 3 and at the blade tips to the rim 13.Between the inner blading 14 and the compressor 2 is a stator blading30.

The assembly of the blades 12 and 14 is illustrated schematically inFIG. 3. Each of the blades 15 of the outer ring 12 is provided with atrapezoidal-section root 16 which is inserted into a longitudinal grooveof corresponding section 17, machined in the radially outer face of therim 13, in order to form a dovetail assembly. Each of the blades 18 ofthe inner ring 14 is provided with a tip 19 presenting lugs 20 spacedaxially and directed radially outwards, and with a root 21 presentinglugs 22 spaced axially and directed radially inwards. The lugs 20 engagebetween and around circular and axially spaced ribs 23 on the radiallyinner face of the rim 13, and are pivotally attached thereto by means ofa pin 24 passing through aligned bores in both lugs and ribs. Thepivotal assembly of the root of each blade 18 is effected in similarmanner by means of a pin 25 passing through circular and axially spacedribs 26 formed at the periphery of the hub 3, and through the lugs 22which are engaged around and between said ribs. The pins 24 and 25 aredisposed parallel to the common rotational axis X X of the hub 3 of therotor 5. The blades 18 are assembled so that they are all inclined tothe hub 3 in the same direction and through the same angle in relationto the circumference, that is to say so that the plane passing throughthe pins 24 and 25 of one and the same blade 18 is tangential to animaginary cylinder 27 of axis X X (see FIG. 4).

From a consideration of FIG. 4, it will be seen that when the hub 3 isrotated in the direction of the arrow F, the blades 18 of the inner ring14 transmit the torque to the rim 13 and are simply subjected to atensile force. Under the effect of the centrifugal forces applied to itby the outer blading 12, the rim l3 expands elastically and thisexpansion is not resisted since the rim 13 isfree to expand due to thepivotal assembly of the blades 18. In other words, considering FIG. 4,it will be seen that the rim l3 expands to occupy the position shown inbroken lines at 13a, the blades 18 pivoting about their axes defined bythe pins 24 and 25 in order to adopt the positions shown in broken linesat 18a. This displacement does not subject the blades 18 to any bendingstress. It should be noted that FIG. 4 is schematic, and the expansionof rim 13 to 13a has been very much magnified for emphasis.

The blades 18 of the inner blading ring 14 are set to give atransparency condition", that is to say are set in order not to produceany aerodynamic effect upon the air entering at 9 (FIG. 1), when theengine is operating in design conditions. This transparency condition isillustrated by FIG. 2 which shows the diagram of the velocities at entryand exit in the case of a blade profile 18 which has a tengentialvelocity U in the direction of the arrow F. The air entering the blading14 at the absolute velocity V has a relative velocity W in relation tothe profiles of the blades 18, which is the resultant of V and -U Thetransparency condition is achieved by inclining the blades 18, inrelation to the direction of the incident air 9, at an angle at whoserelat ive velocity W is inclined v i sa-vis the absolute velocity V Therelative velocity W1 at exit from the blade is equgl to W so that theresultant of t his relative velocity W1 and the tangential velocity Uproduces, at the exit from the blading l4, a n absolute velocity V1equal to the absolute velocity V at the blade entry.-

Because of this transparency condition on the part of the inner blading14 of the fan, this blading consequently has no influence upon theairflow 9 entering the compressor, at any rate in the region of thedesign point. The chief advantage of this arrangement is that the jetengine, from the point of view of regulation, behaves as a twin-spoolengine, that is to say it behaves as if it did not have a fan. The fanserves exclusively to supply the annular duct 11 with the requisiteby-pass air. Moreover, since the by-pass ratio is independent of thecompression ratio of the twin-spool compressor, it is possible veryeasily to modify the engine to obtain a different by-pass ratio, merelyby changing the fan and without interfering with the rest of the engine.This kind of modification might for example be desirable in order toadapt the engine to a STOL aircraft which requires an engine with a highby-pass ratio.

In addition, since the blading 14 produces no aerodynamic effect, theairflow 9 exerts virtually no stress on it. Therefore, the blades 18need only have sufficient mechanical strength to enable them to transmitthe torque of the hub 3 to the rim 13, a function which subjects them totensile stresses since they do not have to withstand any aerodynamicforces.

The present invention is applicable not only to an engine but equally toan industrial gas turbine engine, in which case it provides thefollowing advantages the centrifugal stresses are not transmitted to thecentral hub the blades of the inner blading do not have to withstand anybending stresses the assembly of the inner blading is particularlysimple the mass of the inner blading can be kept very low.

We claim:

1. In a gas turbine unit comprising axial-flow compressor means, a fanfitted with a hub and mounted for rotation coaxially in front of saidaxial compressor means, a rim mechanically connected to said hub torotate therewith, an array of outer blading attached to said rim andextending outwardly thereof means for rotating the fan in order that theouter blading produces an annular outer airflow surrounding thecompressor means, and means for rotating the compressor means in orderto compress an inner airflow passing between the hub and the rim andthus to supply combustion air to the gas turbine unit; and improved modeof mechanical connection of the rim to the hub comprising a plurality ofblades each connected at its tip and root, respectively, to said rim andto said hub, for limited pivotal movement about pivot axes extendingparallel to the axis of rotation of said hub, said blades all beinginclined in the same direction and at the same angle in relation to thehub circumference whereby said pivoted blades can accommodate toexpansion and contraction of said rim under inertial and thermal forces.

2. A gas turbine unit as in claim 1 wherein each of said pivotallymounted blades is oriented at a setting angle selected to correspondwhen the unit is operating under design condition with the relativedirection of said inner airflow considered in relation to said blades,whereby said blades are in a thrust-free condition, with respect to saidairflow under said design conditions.

3. A gas turbine unit as claimed in claim 2, for an aircraft by-passtype gas turbine jet engine, wherein the means for driving saidcompressor means is a turbine and the means for driving said fan is aseparate turbine.

4. A gas turbine unit as claimed in claim 11, wherein said separateturbine is adapted to drive said fan at a speed imparting to the outerblading a peripheral velocity sufficiently low as to produce a very lownoise level, and the outer blading comprises a sufficient number ofblades to achieve in operation a substantial compression ratio.

5. A gas turbine as in claim 4 wherein each of said pivotally mountedblades extends between said hub and said rim at an angle at leastslightly inclined from a radius passing from the axis of fan rotationthrough a opposite to the direction of fan rotation whereby the pivotaxis of such blade. forces acting on said blades during fan rotation aresub- 6. A gas turbine as in claim 5 wherein said blades are stantiallyin tension.

inclined from root to tip at an angle having a direction

1. In a gas turbine unit comprising axial-flow compressor means, a fanfitted with a hub and mounted for rotation coaxially in front of saidaxial compressor means, a rim mechanically connected to said hub torotate therewith, an array of outer blading attached to said rim andextending outwardly thereof; means for rotating the fan in order thatthe outer blading produces an annular outer airflow surrounding thecompressor means, and means for rotating the compressor means in orderto compress an inner airflow passing between the hub and the rim andthus to supply combustion air to the gas turbine unit; and improved modeof mechanical connection of the rim to the hub comprising a plurality ofblades each connected at its tip and root, respectively, to said rim andto said hub, for limited pivotal movement about pivot axes extendingparallel to the axis of rotation of said hub, said blades all beinginclined in the same direction and at the same angle in relation to thehub circumference whereby said pivoted blades can accommodate toexpansion and contraction of said rim under inertial and thermal forces.2. A gas turbine unit as in claim 1 wherein each of said pivotallymounted blades is oriented at a setting angle selected to correspondwhen the unit is operating under design condition with the relativedirection of said inner airflow considered in relation to said blades,whereby said blades are in a thrust-free condition, with respect to saidairflow under said design conditions.
 3. A gas turbine unit as claimedin claim 2, for an aircraft by-pass type gas turbine jet engine, whereinthe means for driving said compressor means is a turbine and the meansfor driving said fan is a separate turbine.
 4. A gas turbine unit asclaimed in claim 3, wherein said separate turbine is adapted to drivesaid fan at a speed imparting to the outer blading a peripheral velocitysufficiently low as to produce a very low noise level, and the outerblading comprises a sufficient number of blades to achieve in operationa substantial compression ratio.
 5. A gas turbine as in claim 4 whereineach of said pivotally mounted blades extends between said hub and saidrim at an angle at least slightly inclined from a radius passing fromthe axis of fan rotation through a pivot axis of such blade.
 6. A gasturbine as in claim 5 wherein said blades are inclined from root to tipat an angle having a direction opposite to the direction of fan rotationwhereby the forces acting on said blades during fan rotation aresubstantially in tension.